Hybrid strapdown guidance system

ABSTRACT

A guidance system for guiding a missile in its boost phase. The system uses a two-axis spaced-fixed platform with pitch and yaw gimbals controlled by a dual-axis gyro that is mounted on the platform. A single-axis roll gyro is strapped-down to the missile frame for controlling the missile in roll. The platform is controlled in roll by the missile frame. The system has three accelerometers on the platform, two being a part of the dual-axis gyro for sensing acceleration in pitch and yaw, and the third being a range accelerometer for sensing acceleration along the longitudinal axis of the missile. The sensed pitch and yaw acceleration voltages are processed in a guidance computer and are compared therein with programmed flight path voltages. If there is any difference between corresponding voltages in pitch and yaw, corrective guidance command signals are generated. The guidance command signals are fed to a pitch-and-yaw signal processor for processing. This processor provides as outputs command signals to airframe actuators for guiding the missile in pitch and yaw. Output signals from the roll gyro are likewise applied, through a roll control signal processor, to the same airframe actuators for guiding the missile in roll. The pitch and vertical acceleration voltages are so processed in the guidance computer that an engine cut-off signal is generated at the proper time.

' United States. Patent 1191 Stripling July 17, 1973 HYBRID STRAPDOWNGUIDANCE SYSTEM [57] ABSTRACT.

[75] Inventor: William W. Stripling, Huntsville, A guidance System forguiding a missile in its boost phase. The system uses a two-axisspaced-fixed platform with pitch and yaw gimbals controlled by a dual-[73] Asslgneei The unltedstates of Amen as axis gyro that is mounted onthe platform. A single-axis represented the Secretary the roll gyro isstrapped-down to the missile frame for con- Army, Washington, DC.trolling the missile in roll. The platform is controlled in 2 Filed; 4197 roll by the missile frame. The system has three accelerometers onthe platform, two being a part of the dual- [21] APP!" 168,883 axis gyrofor sensing acceleration in pitch and yaw, and

the third being a range accelerometer for sensing accel- 521 US. Cl.244/32 eratieh along the longitudinal axis of the missile The [51] Int.Cl. sensed Pitch and y acceleration Voltages are P" [58] Field of Search244/32 cessed in a guidance computer and are p d therein with programmedflight path voltages. If there [56] References m is anydifferencebetween corresponding voltages in UNITED STATES PATENTS pitch and yaw,corrective guidance command signals are generated The guidance commandsignals are fed gz z to a pitch-and-yaw signal processor for processing.This 3057211 10/1962 Duncan Jail I .11.: 244 312 pmcessm pmvides as P P"W e i 3:283:573 11/1966 Bishop e alm 244/32 UX frame actuators forguidmg the missile 1n pitch and 3,547,381 12/1970 Shaw 244 32 OutputSignals from the roll y are likewise P- A ROLL AXIS AW ACCEL PITCH.ACCEL.

V TICAL plied, through a roll control signal processor, to the sameairframe actuators for guiding the missile in roll. The pitch andvertical acceleration voltages are so processed in the guidance computerthat an engine cut-off signal is generated at the proper time.

3 Claims, 4 Drawing Figures flea AXIS 38 i PITCH a SIGNAL PROCESSOR GUIDANCE COMPUTER ROLL AMPLIFIER COMMAND PITCH CONTROL ROLL ANGLE COMMANDSI I P11 gmpgggi I law J com/mun CUTOFF uw cummc coma/mu nou. ANGLE INEPNENIED Jul I mm 3, TAG .2 81

SHEET M UF 4 PITCH YAW GUIDANCE GUIDANCE COMMANDS COMMANDS M fi w M g II l I E? 380 I I ,38c 38e YAW ANGLE E Q1358!) II I l I l PITCH ANGLE I iI ;8f 1 I I I I I I YAW CONTROL I I COMMAND I PITCH CONTROL I I COMMANDL h D h i l FIG. 4

M o/PL HYBRID STRAPDOWN GUIDANCE SYSTEM BACKGROUND OF THE INVENTION Inguiding a ground-to-ground missile during the boost phase of flight, itis desired that the missile arrive at an engine cut-off point with acertain attitude and velocity. The missile then follows a free flighttrajectory to the target. The orientation of the missile during thisboost phase must be controlled in roll, pitch and yaw for arriving atthis engine cutoff point with proper speed and attitude. The missleusually begins its trajectory by rising vertically for a few seconds. Itthen rolls to the proper heading. Once in the proper roll direction, themissile executes its pitch maneuver. The flight of the missile can beconsidered to pitch-over down range in one vertical plane with noacceleration in a traverse direction.

One commonly used means of guiding a missile in such a vertical liftoffboost phase uses a strapdown, i.e., hard-mounted to the missile frame,three-axis inertial system for guiding the missile in roll, pitch, andyaw. Strapdown systems such as this place stringent requirements on thegyro dynamic range performance in maintaining a space fixed reference ina guided missile. A three-axis platform may use threesingle-degree-offreedom gyros arranged'orthogonally in such a way thatthe spin axes of two of the gyros are in parallel planes. Any angularmotion of the platform about the input axis of either of these gyroshaving spin axes in a parallel plane is sensed by the pickofl of theother gyro.

The basic improvement that the present system allows over a conventionalthree-axis platform is the simplification of using the missile frame asa third gimbal for maintaining alignment of the accelerometer in aspecific inertial reference frame. When a two-axis platform having atwo-degree-of-freedom gyro and accelerometer with input axes alignedwith the pitch and yaw axes is augmented by a strapdownsingle-degree-offreedom gyro in the roll axes, a simple three-axiscontrol system is provided that can be made more compact than presentthree-axes control systems.

SUMMARY OF THE INVENTION The present invention is a hybrid two-axisspacefixed platform and strapdown single-axis roll gyro guidance systemcombined for use in guiding a missile during its boost propulsion phaseof operation. The strapdownroll gyro is a single-degree-of-freedom gyrothat has its input axis aligned with the longitudinal axis of themissile for sensing missile roll and having an output to controls foraerodynamic surfaces. These surfaces keep the missile roll-stabilized.The platform is controlled in roll by the missile airframe. The two-axisspace-fixed platform has a two-degree-of-freedom gyro and accelerometerwith its two input axes aligned with the pitch and yaw axes of themissile.

The guidance of the missile is based upon output signals fromthreeaccelerometers, all mounted on the space-fixed platform with theirsensitive axes mutually perpendicular. The output signals from theaccelerometers are applied to a guidance computer. Two of theaccelerometers, whose input axes are oriented in pitch and yaw, are apart of a two-degree-of-freedom gyro and accelerometer. A rangeaccelerometer is positioned separately on the space-fixed platform forsensing vertical acceleration of the missile. Output signals from theaccelerometers are compared with programmed flight path signals in theguidance computer. If there are any differences between the actual andthe programmed pitch and yaw voltages, guidance command signals aregenerated, and are applied to a pitch and yaw signal processor. In thesignal processor, the command signals are compared to the actual pitchand yaw signals as derived from transducers between the space-fixedplatform and the missile airframe. The differences between the comparedsignals are applied to airframe actuators as command signals. Theairframe actuators move aerodynamic control surfaces for guiding themissile according to the desired flight path. The pitch and verticalacceleration voltages are processed in the guidance computer forgenerating an engine cutoff signal at a prescribed velocity when themissile is in a prescribed attitude. The missile is then in a freeflight trajectory to the target.

BRIEF DESCRIPTION OF THE DRAWING FIG. 4 is a schematic block diagram ofone of the signal processors of FIGS. 1 and 2.

DESCRIPTION OF THE PREFERRED EMBODIMENT Referring to FIGS. 1 and 2, atwo-axis space-fixed platform 20 is shown mounted within missile 32.Platform 20 has a two-degree-offreedom gyro and accelerometer 10 and arange accelerometer 12 mounted thereon. Device 10 has sensitive axes 5and 7 (FIG. 2). Axis 5 is positioned parallel with the pitch axis of themissile, and axis 7 is positioned parallel with the yaw axis of themissile. A single-degree-of-freedom gyro 34 is separately strapped-downto the frame of the missile with its input axis IA aligned parallel withthe longitudinal or roll axis 8 of the missile. Platform 20 is thencontrolled in roll by the missile frame. Roll gyro 34 produces voltagesat its output in the conventional man ner. These voltages are amplified:in roll control signal processor 36 to provide roll angle commandsignals. These roll angle command signals are applied to actuators 30that, in turn, move aerodynamic surfaces 40 and 42 to move missile 32 inroll.

A typical two-degree-of-freedom gyro and accelerometer device, such asdevice 10, that may be used in this invention is disclosed on page 97 ofthe June 1968 edition of the periodical Space Aeronautics, published byConover-Mast Publication, Inc., New York City. This article is entitled,New Class of Inertial Sensors Combines Wide Range with Small Size andwas written by W. G. James. When missile 32 is positioned for a verticallaunch, input axes 5 and 7 may be respectively aligned parallel to thepitch and yaw axes of thissile 32. Platform 20, on which device 10 ismounted, remains space-fixed after launch and estabilshes a referenceplane to indicate the space orientation of missile 32 about thisplatform during the Iflight. The gyro outputs from device 10 are appliedto pitch and yaw torquers 24 and 22. through amplifiers l6 and I4,respectively, for maintaining platform 20 space-fixed after missilelaunch. The outputs from the accelerometer of device are analog signalsin pitch and yaw that are applied to guidance computer 18 for comparisonwith programmed flight paths in pitch and yaw within the computer. Thevalue of these analog signals at the outputs of the accelerometers arethe actual acceleration of the missile in pitch and yaw, and if theseoutputs are not identical to the programmed flight path equations, thenerror signals, called pitch and yaw guidance command signals, aregenerated at the output of computer 18. Range accelerometer 12 furtherprovides a vertical acceleration signal for missile 32. This verticalacceration signal is also applied to guidance computer 18. The verticalacceleration, pitch acceleration, and yaw acceleration signals areapplied to guidance computer 18. An engine cut-off signal and pitch andyaw guidance commands are derived in 18.

The elements of computer 18 are shown in FIG. 3 in block diagram form.Guidance computer 18 may be similar to a system shown and described onpages 123-127 in a book entitled Navigation and Guidance in Space, byEdward V. B. Stearns by Prentice-Hall Inc., Copyright 1963. Anotherreference of a computer such as guidance computer 18 is shown anddescribed on page 273 of a book entitled Inertial Guidance", by GeorgeR. Pitman, Jr., by John Wiley & Sons Inc., Copyright 1962. The elementsare described hereinbelow with reference to the blocks shown in FIG. 3.If missile 32 moves laterally along axis 5, an acceleration signal isdetected and a pitch acceleration output signal is produced from device10. This pitch acceleration output signal is applied to a first resolver18a. The output of 18a consists of a sine portion and a cosine portion.The sine portion is applied to one input to integrator 18b, and thecosine portion is applied to one input of integrator 18c. Pitchacceleration signals are applied to resolver 18d. The output of 18d hassine and cosine portions respectively applied to inputs of 180 and 18b.The output of integrator 18b is applied to integrator 18: and multiplier18f. The output of 18b is related to vertical velocity, therefore,the-output of l8e is related to vertical displacement. Multiplier 18famplifies the velocity signal from 18b by a factor dependent ontime-to-go to impact. The outputs from 18f and l8e are summed in summer18g. The output of 18g is subtracted from a constant preset signalprovided by flight program generator 18h in subtractor l8i.

The output of 181' feeds a zero-detector (not shown) which detector cutsoff the missile engine when the output from 181' goesto zero.

The output of 18c is related to pitch velocity. When this output isintegrated in integrator'l8j, an output related to pitch displacement isprovided. This output from l8j is subtracted from a pitch output offlight program 7 generator 18h in subtractor 18k. The variable output of18k provides pitch guidance commands.

For yaw commands, the yaw acceleration signals are merely twiceintegrated in integrators 181 and 18m.

As can be seen in FIG. 3, integrators 18b, and 18c, 18c, l8j, 181 andl8m, multiplier 18f and flight program generator 18h are all timed bytimer 18:1. The

' outputs of 181: can be digital or analog. If digital, multiplier 18fcould include a counter having a preset count at liftoff of the missile,which count would be varied by pulses from l8n. Likewise, the variableoutput of 18k could be controlled in a similar manner. It wouldobviously be necessary to convert the guidance commands into analogsignals.

The command signals are applied to first and second inputs of pitch andyaw control signal processor 38 as shown in FIGS. 1 and 2, and as shownin detail in FIG. 4.

As can be seen in FIG. 4, signal processor 38 includes variableamplifiers 38a and 38b to which the pitch and yaw guidance commands arerespectively applied. The pitch angle and yaw angle transducers (26 and28) of FIGS. 1 and 2 produce variable voltages related to the anglesbetween space-fixed platform 20 and missile airframe 32. These variablevoltage outputs from transducers 28 and 26 represent pitch and yawalignment, respectively, and are fed to third and fourth inputs, ofpitch and yaw control signal processor 38. The third input of 38 goes tovariable amplifier 38c and the fourth input goes to amplifier 38d. Theoutputs from amplifiers 38b and 380 are combined in subtractor 38c toprovide a yaw control command. The outputs of 38a and 38d are combinedin subtractor 38f to provide a pitch control command.

Roll control of the missle is accomplished by amplifying the output ofgyro 34 in amplifier 36, and applying the output of 36 to controlactuator 30.

The control command signals are fed to airframe control actuators 30.The actuators 30 furnish signals in pitch to a pitch aerodynamic controlsurface 42 and signals in yaw to a yaw aerodynamic control surface 40.Only one each of surfaces 40 and 42 are shown, but

there are like surfaces 180 apart on missile 32 surface.

The use of four aerodynamic control surfaces is not critical in thepresent guidance system, since three aerodynamic control surfaces'arefrequently used in guiding a missile. However, by using four airframecontrol acutators and four control surfaces, the pitch, yaw and rollcontrol command signals may be summed directly. The use of threeairframe control actuators and aerodynamic control surfaces wouldrequirethe summing of the sine and cosine components of the pitch, yawand roll control command signals within the three airframe controlactuators.

While a specific embodiment of the invention has been shown anddescribed, other embodimentsmay be obvious to one skilled in the art inlight of this disclosure. An example is that the space-fixed platformcould be oriented such that the range accelerometer would be alignedparallel to the slant range velocity vector, instead of perpendicular-tothe longitudinal axis of the missile as described. It should beemphasized the elements l8 and 38 are considered as old and well knownin the art and their exact contents are not limited to the contents ofFIGS. 3 and 4. The gains of the amplifiers in element 38 would bemanually adjusted prior to launch of the missile, in accordance with theparticular mission of the missile.

I claim:

1. A missile with an airframe and a guidance system comprising: aspace-stable platform in said missile; means mounted on said platformfor sensing pitch, and yaw acceleration and providing pitch and yawacceleration signals; means mounted on said platform for sensingvertical acceleration of said platform and for providing a verticalacceleration signal; a guidance computer having inputs to which saidpitch, yaw, and vertical acceleration signals are applied and havingrespective outputs for a missile engine cut-off signal, for pitchv formand the airframe of said missile for respectively responding to theangles between said platfonn and pitch and yaw axes and for providingrespective pitch and yaw angle signals; a signal processor havingrespective inputs for each of said pitch and yaw guidance signals, andsaid pitch and yaw angle signals, said processor also having respectiveoutputs for pitch and yaw command signals; a sensor mounted on saidmissile airframe for sensing roll of said missile about its longitudinalaxis and for providing a signal related to the roll angle; a pluralityof aerodynamic control surfaces on said missile airframe; and arespective plurality of actuators for said control surfaces, with saidpitch and yaw command signals applied to respective inputs of saidactuators, and with said roll signal applied to inputs of saidactuators.

2. The system of claim 1 wherein said guidance computer includes 'meansto provide a pitch dispacement reference signal and a verticaldisplacement reference signal; means for deriving pitch and verticaldisplacement signals from said pitch and vertical acceleration signals;and means for comparing said reference signals and the derived signalsto provide the engine cut-off and pitch guidance commands; and furthermeans in said computer for deriving from said yaw acceleration signalsyaw displacement signals as yaw guidance command signals.

3. The system of claim 2 wherein said pitch and yaw signal processorincludes means for respectively com; paring said pitch commands and saidpitch angle signals and said yaw guidance commands and said yaw anglesignals to provide respectively said pitch and yaw control commands.

1. A missile with an airframe and a guidance system comprising: aspace-stable platform in said missile; means mounted on said platformfor sensing pitch and yaw acceleration and providing pitch and yawacceleration signals; means mounted on said platform for sensingvertical acceleration of said platform and for providing a verticalacceleration signal; a guidance computer having inputs to which saidpitch, yaw, and vertical acceleration signals are applied and havingrespective outputs for a missile engine cut-off signal, for pitchguidance signals and for yaw guidance signals; first and secondtransducers mounted between said stable platform and the airframe ofsaid missile for respectively responding to the angles between saidplatform and pitch and yaw axes and for providing respective pitch andyaw angle signals; a signal processor having respective inputs for eachof said pitch and yaw guidance signals, and said pitch and yaw anglesignals, said processor also having respective outputs for pitch and yawcommand signals; a sensor mounted on said missile airframe for sensingroll of said missile about its longitudinal axis and for providing asignal related to the roll angle; a plurality of aerodynamic controlsurfaces on said missile airframe; and a respective plurality ofactuators for said control surfaces, with said pitch and yaw commandsignals applied to respective inputs of said actuators, and with saidroll signal applied to inputs of said actuators.
 2. The system of claim1 wherein said guidance computer includes means to provide a pitchdispacement reference signal and a vertical displacement referencesignal; means for deriving pitch and vertical displacement signals fromsaid pitch and vertical acceleration signals; and means for comparingsaid reference signals and the derived signals to provide the enginecut-off and pitch guidance commands; and further means in said computerfor deriving from said yaw acceleration signals yaw displacement signalsas yaw guidance command signals.
 3. The system of claim 2 wherein saidpitch and yaw signal processor includes means for respectively comparingsaid pitch commands and said pitch angle signals and said yaw guidancecommands and said yaw angle signals to provide respectively said pitchand yaw control commands.